Gas turbine engine

ABSTRACT

A gas turbine engine has, in a cross section in the axial direction of an annular inner peripheral wall connected to a blade main body of a turbine blade, a concave part and a convex part on the front edge side and the rear edge side. The concave part has a negative curvature and is concave toward the axis, and the convex part has a positive curvature and is convex away from the axis. The flow rate on the upper face of the blade main body can be reduced in the concave part on the front edge side, thus suppressing generation of a shock wave, and the flow rate can be increased in the convex part on the rear edge side following the concave part, thus smoothly changing the flow rate on the upper face of the blade main body and thereby minimizing the pressure loss. In this way, the thickness of the blade main body can be reduced while ensuring the performance of the gas turbine engine, thereby contributing to a reduction in weight.

This application is the national phase under 35 U.S.C. §371 of PCTInternational Application No. PCT/JP00/09150 which has an Internationalfiling date of Dec. 22, 2000, which designated the United States ofAmerica.

FIELD OF THE INVENTION

The present invention relates to a gas turbine engine in which turbineairfoils are disposed in a radial direction in an annular gas passagedefined by an inner peripheral wall and an outer peripheral wall.

BACKGROUND ART

Japanese Patent Application Laid-open No. 11-241601 discloses anaxial-flow gas turbine engine comprising stator vanes and rotor blades,wherein a cross section in the axial direction of an inner peripheralwall of a casing to which the stator vanes and the rotor blades areconnected has a concave part that recedes radially inward relative to astraight line joining the front edge of the root of the stator vane onthe front side and the rear edge of the root of the rotor blade on therear side.

Furthermore, U.S. Pat. No. 5,466,123 discloses an arrangement in which across section orthogonal to the axial direction of an inner peripheralwall of a casing supporting stator vanes of a gas turbine engine hassinusoidally alternating concave parts and convex parts.

It should be noted here that when the thickness of a turbine airfoil ofa gas turbine engine is decreased in order to reduce the weight of theairfoil without changing the material thereof, it is necessary toincrease its stagger angle γ (see FIG. 1), but since the curvature ofthe front half on the upper face of the airfoil increases with anincrease in the stagger angle γ, the flow rate on the upper face of theairfoil accelerates and decelerates rapidly and, in particular, in thecase of a high load blade having a high rotor inlet relative Machnumber, a shock wave is generated beyond a critical Mach number, and asa result there is the problem that the pressure loss increases.

DISCLOSURE OF THE INVENTION

The present invention has been carried out in view of theabove-mentioned circumstances, and it is an object of the presentinvention to suppress the occurrence of a shock wave when the staggerangle is increased as a result of decreasing the thickness of a turbineairfoil of a gas turbine engine, thereby preventing any increase in thepressure loss.

In order to accomplish the above-mentioned object, in accordance withthe present invention, there is proposed a gas turbine engine comprisingturbine airfoils disposed in a radial direction in an annular gaspassage defined by an inner peripheral wall and an outer peripheralwall, characterized in that a cross section in the axial direction alonga connecting section of the inner peripheral wall or the outerperipheral wall where the wall is connected to the turbine airfoil has aconcave part on a front edge side having a negative curvature relativeto the direction of flow of gas and a convex part on a rear edge sidehaving a positive curvature relative to the direction of flow of gas.

Furthermore, in addition to the above-mentioned arrangement, there isproposed a gas turbine engine wherein the height of the convex part isat most 10% of the length, in the radial direction, of the gas passage.

Moreover, in addition to the above-mentioned arrangement, there isproposed a gas turbine engine wherein the cross section in the axialdirection along the connecting section has at least one point ofinflection between the front edge and the rear edge.

Furthermore, in addition to the above-mentioned arrangement, there isproposed a gas turbine engine wherein, among the at least one point ofinflection, the point of inflection that is the closest to the frontedge side is positioned forward relative to the central position of thechord of the turbine airfoil.

Moreover, in addition to the above-mentioned arrangement, there isproposed a gas turbine engine wherein the absolute value of the negativecurvature of the concave part is smaller than the absolute value of thepositive curvature of the convex part.

Furthermore, in addition to the above-mentioned arrangement, there isproposed a gas turbine engine wherein the axial position of the concavepart is set so that the axial position of a minimum negative pressurepoint that is the closest to the front edge of the turbine airfoilconnected to a flat connecting section is present within the range ofthe concave part.

Moreover, in addition to the above-mentioned arrangement, there isproposed a gas turbine engine wherein the front end of the concave partis positioned to the rear of the front edge.

When the thickness of the turbine airfoil of the gas turbine engine isdecreased in order to reduce the weight, the stagger angle requiredincreases, the flow rate of combustion gas on the upper face of thefront half of a blade main body accelerates and decelerates rapidly and,in particular, in the case of a high load blade having a high rotorinlet relative Mach number, the flow rate reaches a critical Machnumber, thus generating a shock wave and thereby causing a largepressure loss and degrading the performance of the gas turbine engine.However, in accordance with the present invention, since the crosssection in the axial direction along the connecting section of the innerperipheral wall or the outer peripheral wall of the gas turbine enginewhere the wall is connected to the turbine airfoil has a concave part onthe front edge side and a convex part on the rear edge side via a pointof inflection, the concave part having a negative curvature relative toa direction of flow of gas, and the convex part having a positivecurvature, the flow rate on the upper face of the blade main body can bereduced in the concave part on the front edge side, thus suppressinggeneration of a shock wave, and the flow rate can be increased in theconvex part on the rear edge side following the concave part, thussmoothly changing the flow rate on the upper face of the blade main bodyand thereby minimizing the pressure loss. In this way, the thickness ofthe blade main body can be reduced while maintaining the performance ofthe gas turbine engine, thereby contributing to a reduction in weight.

This effect can be exhibited even more strongly by making the height ofthe convex part at most 10% of the length, in the radial direction, ofthe gas passage, positioning the point of inflection between the concavepart and the convex part so as to be forward of the central position ofthe chord, making the absolute value of the negative curvature of theconcave part smaller than the absolute value of the positive curvatureof the convex part, arranging for the minimum negative pressure pointthat is the closest to the front edge of the conventional turbineairfoil to be present within the range of the concave part, andpositioning the front end of the concave part so as to be to the rear ofthe front edge.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 and FIG. 2 show one embodiment of the present invention.

FIG. 1 is a diagram showing the shape of a turbine blade of a gasturbine engine.

FIG. 2 is a diagram showing the shape of an inner wall face along thechord of the turbine blade, the curvature of the inner wall face, andspeed distribution on the blade face.

BEST MODE FOR CARRYING OUT THE INVENTION

A mode for carrying out the present invention is explained below byreference to an embodiment of the present invention illustrated inattached drawings.

FIG. 1 and FIG. 2 show one embodiment of the present invention.

FIG. 1 shows a turbine blade 11 of an axial-flow gas turbine engine, andthe turbine blade 11 is formed from a blade main body 12 positionedoutward in the radial direction, a blade end wall 13 positioned inwardin the radial direction relative to the blade main body 12, and a blademounting part 14 positioned inward in the radial direction relative tothe blade end wall 13. The blade shape of the root part (a partadjoining the blade end wall 13) of the blade main body 12 shown as thecross section X—X in FIG. 1 comprises a front edge 12 a, a rear edge 12b, an upper face 12 d, and a lower face 12 e, and a straight linejoining the front edge 12 a and the rear edge 12 b has a comparativelylarge stagger angle γ relative to the direction of the axis A of the gasturbine engine.

The stagger angle γ of the blade main body 12 of this embodiment is setso as to be large compared with a conventional stagger angle γ of 0° to20°. Setting the stagger angle γ so as to be large compared with theconventional stagger angle in this way makes it possible for the bladethickness of the blade main body 12 to be thin, and as a result theweight of the turbine blade 11 can be reduced by 20% relative to theconventional turbine blade without changing the material.

A tip 12 c on the radially outer end of the blade main body 12 faces anannular outer peripheral wall 15 a of an outer casing 15 with a slighttip clearance 16. An annular hub 17 a is formed on the outercircumference of a blade disc 17 supported rotatably around the axis Aof the gas turbine engine, and a large number of the blade mountingparts 14 of the turbine blades 11 are mounted radially on the hub 17 a.In order to withstand a large centrifugal force acting on the turbineblade 11, the blade mounting part 14 has a plurality of alternatingridges 14 a and grooves 14 b extending in the direction of the axis A ofthe gas turbine engine, and these ridges 14 a and grooves 14 b interlockwith the hub 17 a via concavo-convex engagement.

When a large number of the turbine blades 11 are mounted on the hub 17 aof the blade disc 17, the blade end walls 13 of the turbine blades 11extend integrally in the circumferential direction, thus forming anannular inner peripheral wall 13 a. An annular gas passage 18 is formedbetween the outer peripheral wall 15 a and the inner peripheral wall 13a, and the turbine blades 11 are disposed within the gas passage 18through which combustion gas flows in the direction of the arrow F.Stator vanes, which are not illustrated, are disposed on the front sideand the rear side of the turbine blades 11 in the axial direction.

As is clear from FIG. 2, a part of a cross section in the axialdirection of the annular inner peripheral wall 13 a, which is formedfrom the blade end wall 13 of the turbine blade 11, is formed from acurve. That is, the cross section in the axial direction of the innerperipheral wall 13 a includes, from the front edge 12 a side to the rearedge 12 b side, a first straight line part 19, a first concave part 20,a convex part 21, a second concave part 22, and a second straight linepart 23. The first concave part 20 and the second concave part 22 havenegative curvatures and are concave toward the axis A, and the convexpart 21 has a positive curvature and is convex away from the axis A. Afirst point of inflection a is present in a part where the curvaturechanges from negative to positive, and a second point of inflection b ispresent in a part where the curvature changes from positive to negative.The curvature on the upper face 12 d of the blade main body 12 ispositive in the whole region from the front edge 12 a to the rear edge12 b.

Characteristic features in the cross section in the axial direction ofthe inner peripheral wall 13 a in the present embodiment are that thefirst concave part 20 and the convex part 21 are positioned continuouslyto the rear of the first straight line part 19 following the front edge12 a, and that a minimum negative pressure point that is the closest tothe front side of the conventional blade main body, which has a flatinner peripheral wall 13 a on which the first concave part 20, theconvex part 21, and the second concave part 22 are not formed, would bepresent within the range of the first concave part 20 (the range from apoint d at the front end to the point a at the rear end). The deepestpoint c of the first concave part 20 (at which the distance from astraight line joining the front end d and the rear end a of the firstconcave part 20 becomes a maximum) is desirably positioned in thevicinity of the above-mentioned minimum negative pressure point.Moreover, the first inflection point a is positioned forward of the 50%position of the chord (the intermediate position between the front edge12 a and the rear edge 12 b), and the absolute value of the negativecurvature of the first concave part 20 is set so as to be smaller thanthe absolute value of the positive curvature of the convex part 21. Itis appropriate for the height of the convex part 21 to be at most 10% ofthe radial length of the gas passage 18, that is, the distance betweenthe inner peripheral wall 13 a and the outer peripheral wall 15 a.

It should be noted here that when the stagger angle γ is increased byreducing the thickness of the blade main body 12 in order to decreasethe weight of the turbine blade 11, as shown by the broken line in thegraph of the speed distribution on the blade upper face 12 d in FIG. 2,the speed distribution of combustion gas on the upper face 12 d of theblade main body 12 rapidly increases and then rapidly decreases, therebygenerating a large pressure loss.

However, in the present embodiment, since in the cross section in theaxial direction of the inner peripheral wall 13 a of the blade end wall13, the first concave part 20 and the convex part 21 are continuous, theflow of the combustion gas can be diffused in the radial direction inthe first concave part 20, thus suppressing a rapid increase in the flowrate and thereby preventing the generation of a shock wave. Furthermore,the flow rate of the combustion gas is increased in the convex part 21following the first concave part 20, as shown by a solid line in thegraph of the speed distribution on the blade upper face 12 d in FIG. 2,and the speed distribution of the combustion gas on the upper face 12 dof the blade main body 12 can be increased smoothly, thereby decreasingthe pressure loss.

In this way, by changing only the shape of the cross section in theaxial direction of the inner peripheral wall 13 a of the blade end wall13 of the turbine blade 11, rapid changes in the speed distribution onthe upper face 12 d of the blade main body 12 can be suppressed evenwhen increasing the stagger angle γ, thereby contributing to a reductionin weight by decreasing the thickness of the blade main body 12 whileensuring the performance of the gas turbine engine by minimizing thepressure loss.

An embodiment of the present invention has been explained above, but thepresent invention can be modified in a variety of ways without departingfrom the spirit and scope thereof.

For example, the turbine blade 11 is illustrated as a turbine airfoil inthe embodiment, but the present invention can be applied to a statorvane of a gas turbine engine in the same manner. In this case, thepresent invention can be applied to either or both of an innerperipheral wall connected to the radially inner end of the stator vaneand an outer peripheral wall connected to the radially outer end of thestator vane.

Furthermore, as shown by the solid line in the graph showing the speeddistribution on the blade upper face 12 d in FIG. 2, a rapid change isobserved in the flow rate of combustion gas in the vicinity of the 70%position of the chord, but it is also possible to further reduce thepressure loss by increasing the absolute value of the negative curvatureof the second concave part 22 and enlarging the range of the secondconvex part 22 toward the rear edge, thus smoothing the change in theflow rate in the vicinity of the 70% position of the chord.

Industrial Applicability

The present invention can be applied to an axial-flow gas turbine enginefor an airplane, for stationary use, and for any other purpose.

1. A gas turbine engine comprising turbine airfoils disposed around an axis of the engine in a radial direction in an annular gas passage defined by an inner peripheral wall and an outer peripheral wall, wherein a cross section taken along a plane passing through the axis of the engine, of a connecting section of the inner peripheral wall or the outer peripheral wall where the wall is connected to the turbine airfoil has a concave part on a front edge side having a negative curvature which is concave towards the axis of the engine and a convex part on a rear edge side having a positive curvature which is convex towards the axis of the engine.
 2. The gas turbine engine according to claim 1, wherein the height of the convex part is at most 10% of the length, in the radial direction, of the gas passage.
 3. The gas turbine engine according to claim 1, wherein the cross section of the connecting section has at least one point of inflection between the front edge and the rear edge.
 4. The gas turbine engine according to claim 3, wherein, among said at least one point of inflection, the point of inflection that is the closest to the front edge side is positioned forward relative to the central position of the chord of the turbine airfoil.
 5. The gas turbine engine according to claim 1, wherein the absolute value of the negative curvature of the concave part is smaller than the absolute value of the positive curvature of the convex part.
 6. The gas turbine engine according to claim 1, wherein the axial position of the concave part is set so that the axial position of a minimum negative pressure point that is the closest to the front edge of the turbine airfoil which is connected to a flat connecting section is present within the range of the concave part.
 7. The gas turbine engine according to claim 1, wherein the front end of the concave part is positioned to the rear of the front edge. 